Reversible space plane

ABSTRACT

A reversible aerospace plane includes an air intake at a first end of the aerospace plane, at least one heat exchanger disposed in the aerospace plane, and an engine at a second end of the aerospace plane, wherein the aerospace plane is configured to accelerate in a first direction and configured to glide and land in a second direction, wherein the second direction is substantially in a reverse direction from the first direction.

RELATED APPLICATIONS

This application is a continuation of U.S. non-provisional patentapplication Ser. No. 12/154,185, filed May 19, 2008, issued as U.S. Pat.No. 8,215,589, which is a Continuation-In-Part of U.S. application Ser.No. 12/072,317, filed Feb. 25, 2008, now U.S. Pat. No. 7,690,601, whichis a continuation of U.S. application Ser. No. 11/040,170, filed Jan.21, 2005, now U.S. Pat. No. 7,344,111, which claimed priority from U.S.Provisional Application No. 60/538,417 filed on Jan. 23, 2004.

GOVERNMENT FUNDING

No government funding was utilized for this invention.

BACKGROUND

The present application relates generally to space planes. Moreparticularly, the invention concerns a novel reversible space plane.

The loss of the space shuttle Columbia in 2003 highlights a need for asafer reusable single-stage-to-orbit (“SSTO”). The Columbia included apayload during re-entry, which was not typical for such re-entries. Inaddition to the mass of the payload, problems with the tiled heat shieldled to the catastrophic loss of the Columbia. Due to the shuttle'srelatively small footprint, structural weight, and rapid decent into theatmosphere, it dissipates most of the kinetic energy of orbital velocityin the denser atmosphere, relying exclusively on the heat shield toremain intact. Because of the need to clear the atmosphere relativelyquickly and reliance on boosters, the NASA space shuttle evolved into adaunting behemoth that is very costly to assemble and launch.

U.S. Pat. No. 5,191,761 (“the '761 patent”), owned by the applicant forthe present invention, discloses an air breathing aerospace engine. Thatpatent is incorporated by reference in its entirety. The engine includesa frontal core that houses an oxygen liquefaction system that capturesambient air and liquefies and separates the oxygen. The oxygen may thenbe used in the rocket engine.

U.S. Pat. No. 6,213,431 (“the '431 patent”) owned by the applicant forthe present invention, discloses an aerospike engine. That patent isincorporated by reference in its entirety. An aerospike engine may havea tapered body with a slanted or curved reaction plane. A fuel injectordirects fuel down the reaction plane. The combustion of the fuel on thereaction plane creates a propulsive force across the reaction plane.

What is needed, therefore, is a reversible re-usable SSTO vehicle thatmay be expediently launched to service the rapidly expanding spaceenterprise. A reduction in cost and an improvement in payload capacityare also desires of this growing industry.

BRIEF SUMMARY OF THE INVENTION

In one aspect, the invention relates to a reversible aerospace planethat includes an air intake at a first end of the aerospace plane, atleast one heat exchanger disposed in the aerospace plane, and an engineat a second end of the aerospace plane, wherein the aerospace plane isconfigured to accelerate in a first direction and configured to glideand land in a second direction, wherein the second direction issubstantially in a reverse direction from the first direction.

In another aspect, the invention relates to a method of flying anaerospace plane that includes accelerating to an orbital velocity in afirst direction, re-orienting the aerospace plane, and re-entering anatmosphere in a second direction, wherein the second direction issubstantially in an opposite direction from the first direction.

Other aspects and advantages of the invention will be apparent from thefollowing description and the appended claims.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a view of an ideal sphere moving at supersonic velocity.

FIG. 2 shows a cross section of an embodiment of an aerospace plane inaccordance with one embodiment of the invention.

FIG. 3 shows a schematic of a cooling/condensing system in accordancewith one embodiment of the invention.

FIG. 4A shows one embodiment of an aerospace plane in accordance withone embodiment of the invention.

FIG. 4B shows one embodiment of an aerospace plane with boosters inaccordance with one embodiment of the invention.

FIG. 4C shows one embodiment of an aerospace plane during re-entry, inaccordance with one embodiment of the invention.

FIG. 4D shows one embodiment of an aerospace plane prepared to land inaccordance with one embodiment of the invention.

FIG. 5 shows one embodiment of an aerospace plane with an aerospikeengine in accordance with one embodiment of the invention.

FIG. 6 shows one embodiment of an aerospace plane in accordance with oneembodiment of the invention.

FIG. 7 shows an alternate embodiment of the aerospace plane of theinvention.

FIG. 7A is a generally schematic flow diagram illustrating the air flowthrough the various components of the aerospace plane illustrated in theupper portion of FIG. 7.

FIG. 7B is a generally schematic flow diagram illustrating the air flowthrough the various components of yet another form of the aerospaceplane of the invention.

FIG. 8 is a top plan view of the aerospace plane illustrated in FIG. 7B.

FIG. 9 is a generally perspective view of still another alternate formof aerospace plane of the invention.

DETAILED DESCRIPTION OF THE INVENTION

An aerospace plane in accordance with one or more embodiments of theinvention may be a reversible aerospace plane. The aerospace plane mayinclude an air liquefaction system that enables the aerospace plane totravel at hypersonic velocities in the atmosphere with reduced drag.When operated in the reverse direction, the aerospace plane may exhibita larger drag so that the kinetic energy from an orbital velocity may bedissipated at a higher altitude and over a longer time period thanconventional vehicles.

FIG. 1 shows the ideal situation of a sphere 100 moving at hypersonicvelocity through the atmosphere. The surface 102 of the front half ofthe sphere 100 is an ideal condenser that will condense the incident airto a liquid upon contact with the surface 102. In this ideal model,instead of creating a shockwave in the atmosphere, the sphere 100condenses the air, thereby creating a partial vacuum in front of thesphere 100. The effect of this condensation of air is to reduce the dragexperienced by the sphere 100 to zero. The sphere can move at anunlimited speed through the atmosphere, without creating drag. Thisprinciple may be applied to the present invention to enable hypersonicvelocity at previously unattainable speeds.

FIG. 2 shows a cross section of an aerospace plane 200 in accordancewith one embodiment of the invention. The aerospace plane includes anose cone 201 at the front and a bell rocket engine 202 at the rear. Anair intake 204 allows air to flow into the aerospace plane 200 as itmoves through the atmosphere. The air enters a heat exchanger portion205 of the aerospace plane 200.

FIG. 3 is a schematic of a heat exchanger system 300 in accordance withone embodiment of the invention. Reference may be made to FIG. 2 aswell, to show the respective location of the components in thisparticular embodiment.

Air that is incident on the nose cone 301 (201 in FIG. 2) is cooled by acooling fluid in the nose cone 301. For a conventional aircraft, the airthat is incident upon the front of the aircraft as it moves through theatmosphere is compressed adiabatically. That is, the compression occurswithout substantial heat transfer. As a result, the incident airincreases in temperature. By cooling the air that is incident on thenose cone 301, the compression may be an isothermal compression. Thatis, heat is absorbed from the incident air so that it is compressedwithout a significant increase in temperature.

Generally, isothermal compression requires less energy that a similaradiabatic compression. Because of the lower energy requirement, there isless drag on the aerospace plane (200 in FIG. 2) as it travels throughthe atmosphere.

It is noted, however, that in practice, the incident air may experiencean increase in temperature. For example, incident air, which may have atemperature close to 0 degrees F. at altitude, may be heated to over1,500.degrees F. because of drag for a conventional aircraft travelingat about Mach 5. Precooling the air, as will be described, may result inthe incident air being heated to only 500.degree.F. Thus, thecompression process with pre-cooling more closely approaches theisothermal ideal.

The precooling of the air before it flows through the intake (204 inFIG. 2) may be done using nitrogen gas (or liquid) separated from theincident air, as will be described. A heat exchanger 350 in the nosecone 301 may be used to precool the air.

Upon flowing into the air intake 304 (204 in FIG. 2), the incident airenters a heat exchanger/condenser portion (205 in FIG. 2) of theaerospace plane. In the embodiment shown in FIG. 3, the incident air iscooled and condensed in three stages, 310, 320, 330. More or less thanthree stages may be used without departing from the scope of the presentinvention.

An aerospace plane in accordance with the invention may include ahydrogen tank 341 for storing an amount of hydrogen that is necessaryfor propulsion. The hydrogen it typically stored in liquid form, andtherefore, must be kept below −423.degree.F., the boiling point ofhydrogen. This liquid hydrogen must be evaporated before it may be usedas a propellant in the engine 302. To evaporate the hydrogen, it isconvenient to flow the hydrogen through heat exchangers (e.g., 310, 330in FIG. 3) so that cooling an condensing of the incident air may beaccomplished at the same time.

As shown in FIG. 3, hydrogen from the hydrogen storage tank 341 ispumped through the third stage heat exchanger 330, where the lowtemperature and the heat of vaporization are used to condense theincident. As will be discussed later, in some embodiments, only theoxygen from the incident air is liquefied.

Hydrogen has a specific heat of 3.425 BTU/lb-.degree.F. and a heat ofvaporization of 191.7 BTU/lb. Oxygen, on the other hand, has a specificheat of 0.219 BTU/lb-.degree.F. and a heat of vaporization of 91.7BTU/lb. The greater values for hydrogen provide an advantage in coolingand condensing the oxygen.

Following the third stage, the hydrogen, typically in gaseous form,flows to the first stage 310, where it is used to continue the coolingprocess of the incident air following precooling from the nose cone 301.The hydrogen may then be pumped to the engine for use as a propellant orfuel.

The incident air, following the precooling at the nose cone 301, flowsthrough the air intake 304 (204 in FIG. 2) and into the first stage heatexchanger 310. In the first stage 310, the air cooled, and the energyfrom the air is used to heat the hydrogen to an appropriate temperaturefor combustion in the engine.

The cooling of the incident air continues in the second stage heatexchanger 320. In the embodiment shown in FIG. 3, the coolant in thesecond stage 320 is liquid oxygen, which may be from an oxygen tank 345or it may be the liquefied oxygen that results from the condensation ofthe oxygen in the air in the third stage 330.

In the third stage 330, at least a portion of the oxygen in the air isliquefied by giving up energy to the liquid hydrogen coolant in thethird stage 330. Air is mostly comprised of oxygen (about 20%) andnitrogen (about 80%). The boiling point of oxygen (i.e., thetemperature, at 1 ATM, below which oxygen is a liquid) is −180.degree.F.and the boiling point of nitrogen is −230.degree.F. This differenceenables the condensation of some or all of the oxygen in the incidentair, without liquefying any of the nitrogen in the incident air.

It is noted that the invention does not preclude the liquefaction ofnitrogen in the incident air. However, there may be certain advantagesto liquefying only the oxygen in the incident air. For example, coolingpotential needed to liquefy the nitrogen may be saved and used for otherpurposes, such as tanking additional oxygen. Also, liquefying thenitrogen in the incident air would require larger and more massive heatexchangers, which may adversely affect the available payload. Inaddition, the cooled nitrogen gas may be used for cooling purposes, aswill be described.

Following the third stage 330, the incident air may be separated into anoxygen component and a nitrogen component. The nitrogen component, shownat 323, may flow to the precooler heat exchanger 350 in the nose cone301 of the aerospace plane. The oxygen component may flow to the secondstage heat exchanger 320, where in may be evaporated into gas for use inthe engine 302. Additionally, the liquid oxygen from the incident airmay be pumped to a storage tank 345 for storage and later use—forexample, it may stored for use in space, where there is no atmosphere toprovide incident air.

Liquefying oxygen from the atmosphere during flight presents numerousadvantages. First, collecting and liquefying oxygen during flightgreatly reduces the amount of tanked liquid oxygen that must be storedon-board before lift off. A non-air breathing rocket must carry all ofthe oxygen that will be used during the entire flight. This represents asignificant mass. The hydrogen combustion reaction with oxygen requires2 moles of hydrogen for every mole of oxygen (H.sub.20 has two hydrogenatoms for every atom of oxygen). But because oxygen is 16 times heavierthan hydrogen, the required oxygen has 8 times the mass of the requiredhydrogen. In an air-breathing rocket, the oxygen may be distilled fromthe atmosphere, thus saving a substantial amount of mass.

Appendix A to this application includes two tables showing the amount ofpre-launch mass, including fuel and oxygen, that is required to propelone pound of payload into orbit. The fuel in this case is hydrogen. Thetwo cases are for a non-air breathing aerospace plane and an airbreathing aerospace plane. Starting with an orbital velocity of 25,000ft./sec, the chart shows calculations working backwards to zerovelocity. In each step, the difference in kinetic energy (DKe) is usedto determine the differential masses of the fuel (DH2, DO2) required toachieve the kinetic energy differential. The masses are thencumulatively added to the mass (MM) of the rocket.

The upper chart shows that for a non-air breathing rocket, 9.116 poundsof takeoff weight are required to get 1.000 pounds of payload to anorbital velocity of 25,000 ft./sec. The lower chart represents an airbreathing rocket. At velocities below 14,000 ft./sec, which representflight in the atmosphere, the differential in oxygen mass (DO2) is zero.This is because the oxygen may be condensed from the atmosphere, asdescribed above. The lower chart shows that only 5.183 pounds of takeoffweight is needed to propel 1.000 pound of payload to an orbital velocityof 25,000 ft./sec. For embodiments where a fraction of the liquefiedoxygen is tanked for later use, the required takeoff weight may be evenlower.

Appendix B shows similar charts for a rocket fueled with methane. Anon-air breathing rocket may require 23.941 pounds of takeoff weight topropel 1.000 pound of payload to an orbital velocity of 25,000 ft./sec,where an air breathing rocket may require only 10.572 pounds of takeoffweight.

It is further noted that a hydrogen slush may be tanked instead ofsimple liquid hydrogen. A slush includes partially frozen hydrogen thatis still able to be pumped. This would increase the cooling capacity ofan aerospace plane by as much as 13%, resulting in a payload increase ofas much as 10%.

FIG. 4A shows a reversible aerospace plane 400 in accordance with oneembodiment of the invention. A reversible aerospace plane is one that iscapable of takeoff/acceleration in one direction, butdeceleration/re-entry and landing in a reverse direction.

The aerospace plane 400 includes a nose cone 401, and air intake 404,and a conventional bell rocket engine 402. In addition, the body of theaerospace plane 400 includes two wings 411? 412. During anacceleration/takeoff mode, the aerospace plane 400 may be propelled bythe engine 402 in the direction shown by the arrow 405. In thisdirection, the wings 411, 412 form a “hyper foil,” which is used to meanthat they present a small profile to the incident air, and the drag isminimized. The wings 411, 412 may form an air foil so that they willprovide lift during atmospheric flight. In addition, lift may begenerated by the angle of attack of the aerospace plane 400.

The nose cone 401 and the associated heat exchangers (e.g., 350 in FIG.3) may be constructed of a light and relatively inexpensive material sothat the nose cone 401 may be jettisoned from the aerospace plane 400before re-entry. During re-entry, the aerospace plane 400 may fly in anopposite direction, and the nose cone would no longer be needed. Theconstraints of heat exchanger design may require that the nose cone 401be formed in such a way that it would not be able to withstand theforces and heat of re-entry. In addition, a nose cone may present ahazard or obstruction during landing. Thus, it may be jettisoned fromthe aerospace plane 400, as will be explained.

The aerospace plane 400 in FIG. 4A may be used with a piggy-backarrangement to gain an initial altitude and airspeed. For example, alarger plane may be used to carry the aerospace plane 400 from theground to an altitude of 30,000 ft.-50,000 ft. From this point, the bellengine 402 may be engaged to provide the thrust to achieve orbit.

FIG. 4B shows an aerospace plane 420 with solid rocket boosters 421,422, similar to the boosters that have been used with the NASA spaceshuttle orbiter. The boosters 421, 422 may be used to provide lowaltitude thrust for the aerospace plane 420. The boosters 421, 422 maybe jettisoned once they have been spent.

FIG. 4C shows one embodiment of a reversible aerospace plane 430 duringa re-entry phase. The aerospace plane 430 is flying in a reverseorientation from the aerospace planes 400, (420 shown in FIGS. 4A and4B). This may be accomplished by simply using orientation thrusters torotate the aerospace plane 430 180.degree. while in orbit and beforere-entry begins. In FIG. 4C, the nose cone (401 in FIG. 4A) has beenjettisoned. In addition, at the forward section of the aerospace plane430 in this mode, the engine (402 in FIG. 4A) has been likewisejettisoned for aerodynamic and control purposes.

The aerospace plane 430 and its wings 431, 432 are formed so that in thereverse direction, they create a “para foil”, that is, they are formedto have rounded edges that present a large profile and create more dragthat when the aerospace plane 430 flies in the takeoff direction (e.g.,the direction shown in FIG. 4A). As shown in FIG. 4C, the aerospaceplane 430 may be pitched upwardly so as to create even more drag.

The drag on the aerospace plane 430 in the reverse direction enables theaerospace plane 430 to dissipate a large amount of kinetic energy in theupper atmosphere, where atmospheric density it low enough that theaerospace plane 430 will not generate temperatures that requiresophisticated heat shielding.

For example, the NASA space shuttles will generally re-enter the denseatmosphere at very high speeds. The space shuttle will slow to normalair velocities within about a quarter of a full orbit. For example, whenlanding in Florida, it is typical for a space shuttle to begin slowingdown at a position near Hawaii. The shuttle will then slow down and landin the distance between Hawaii and Florida.

An aerospace plane 430 in accordance with the invention may have asufficient drag so that slowing down may be accomplished at a muchhigher altitude and over a longer distance. For example, an aerospaceplane 430 may slow from orbital velocity over two complete orbits aroundthe Earth, taking a much longer time. The additional time enables theaerospace plane 430 to dissipate the heat associated with slowing downso that sophisticated heat shielding is not required. Further, thestructure and required propellant of such an aerospace plane may enableit to be substantially lighter than previous vehicles. A reduction inmass will also reduce the kinetic energy that must be dissipated duringre-entry.

It is noted that an aerospace plane in accordance with the invention maybe referred to a traveling in a “reverse direction.” In practice, anaerospace plane may be oriented in a reverse situation, even though thevector of travel for the aerospace plane has not itself reversed. Theuse of “reverse direction” is meant to indicate a reverse orientation ofthe aerospace plane.

FIG. 4D shows the aerospace plane 430 in a maneuvering/landing mode. Theaerospace plane 430 is pitched downward for gliding, maneuvering, andlanding. The wings 431, 432 may form an airfoil to generate lift thatwill aid in the maneuverability of the aerospace plane 430.

It is also noted that an aerospace plane in accordance with theinvention may be manned or unmanned. A remotely controlled aerospaceplane may be used while still gaining the advantages of the presentinvention. A manned aerospace plane is also within the scope of theinvention. The reduced temperatures during re-entry provide asignificantly safer re-entry phase than with the existing space shuttledesign.

FIG. 5 shows another embodiment of an aerospace plane 500 in accordancewith the invention. The aerospace plane 500 includes a nose cone, andair intake 504, and wings 511, 512, as the embodiment shown in FIG. 4A.The illustrated difference is that the aerospace plane 500 in FIG. 5includes an aerospike engine 502 instead of a bell nozzle. An asonicaerospike engine is disclosed in U.S. Pat. No. 6,213,413 (“the '413patent”), which is owned by the applicant for the present invention.That patent is incorporated by reference in its entirety.

The aerospike engine 502 shown in FIG. 5 includes a primary reactionplane 521, and two secondary reaction planes 522, 523. Any arrangementof reaction planes may be devised for an aerospike engine withoutdeparting from the scope of the invention.

As disclosed in the '413 patent, an aerospike engine is able to operatemore efficiently than a bell nozzle at a variety of altitudes. Becauseof this feature, an aerospace plane 500 with an aerospike engine 502 maybe able to takeoff on a runway, using the thrust from only the aerospikeengine. In this regard, an aerospace plane 500 forms a self-sufficientSSTO vehicle that may takeoff from a runway, achieve an orbitalvelocity, orbit the Earth, re-enter the Earth's atmosphere in a reversedirection, and land. Advantageously, such a aerospace plane 500 may notrequire the use of boosters or a piggy-back.

FIG. 6 shows another embodiment of an aerospace plane 600 in accordancewith the invention. The aerospace plane 600 does not include a nosecone. Instead, the entire aerospace plane forms a wing-type structure,and there is an air intake 604 at a first end of the aerospace plane600. An engine 602 is located at the other end, and in the embodimentshown in FIG. 6, the engine 602 is an aerospike engine. The aerospaceplane 600 is shown with boosters 611, 612 that may be jettisoned. Insome embodiments, and aerospace plane 600 does not include boosters. Forexample, an aerospace plane 600 may include an aerospike engine 602 thatenables the aerospace plane 600 to takeoff, fly to orbit, and landwithout the need for boosters. Additionally, a piggyback may be used.

In a takeoff/acceleration mode, the aerospace plane 600 travels in afirst direction 605. Incident air flows into the air intake 604, and isthen cooled and condensed, thereby reducing the drag on the aerospaceplane 600 at hypersonic velocity. The engine 602 may be used to propelthe aerospace plane 600. Upon reaching orbital velocity, the air intake604 may be closed.

For a re-entry/deceleration/landing mode, the aerospace plane 600 maytravel in a reverse direction 606. The engine, which may be a bellnozzle in some embodiments, may be jettisoned. An aerospike engine mayadapted to withstand the forces and temperatures of re-entry, or anaerospike engine may be retracted for re-entry.

Turning now to FIGS. 7 and 7A of the drawings, yet another embodiment ofthe reversible aerospace plane of the invention is there shown ingenerally designated by the numeral 700. This latest embodiment of theinvention is similar in many respects to the previously describedembodiments, but includes several novel features not illustrated anddescribed in connection with the embodiments of FIGS. 1 through 6.

As in the earlier described embodiments, the nosecone 701 of this latestform of aerospace plane forms the leading edge of the aerospace plane.The nosecone is configured so as to act both as a shock cone and heatexchanger and, in a manner presently to be described, is controllablychilled. Nosecone 701 may be finned, warped or dimpled to augment theheat transfer capability both inside and outside the nose cone.Alternatively, nosecone 701 may be pointed or rounded and may also beelectroplated, etched, or maybe of a bi-metallic construction. Asbefore, nose cone 701 is configured to be jettisoned before a re-entry.

The organpipe 702 of this latest form of the invention comprises acontinuation of the nosecone 701 and allows the precooled ambient air703 to enter the intake aperture 704. The organpipe tubes may be ofconstant diameter or they may be diverging in area and function to allowentrainment of the intake air 703 into the plenum 704, while at the sametime maintaining continued isothermal compression of the intake air 703.The organpipe tubes may be finned, warped, dimpled, electroplated,etched, or they'd be of a bimetallic construction.

As in the earlier described embodiments of the invention, the intake air103 constitutes the primary source of oxygen (propellant) driving therocket engine thru the hypersonic regime. The intake air is isothermallycompressed by super-cooling of the shock front via the nosecone andorgan pipe heat exchangers 701 and 702 respectively. In the absence of achilled/super-cooling nosecone the intake air will be adiabaticallycompressed via the incipient shock wave. Adiabatic compression resultsin searing high temperatures (and ultimate dissociation of the rarifiedair at high altitudes) due to the trapped heat in the shock front. Inthe limit of perfection, super-cooling will morph the intake air into(cool) isothermal compression. Analytically isothermal compressionconstitutes infinitesimal small (adiabatic) compression steps at aconstant temperature. In the limit of the compression>>zero(infinitesimal small compression steps), adiabatic compression morphsisothermal compression. Super-cooling hence facilitates the desiredcompression regression. Because of the volumetric bulk of the rarifiedair at high altitudes, additional compression is essential to mitigatethe size/efficacy of the liquefaction plant.

The intake plenum 704 of this latest form of the invention funnels theisothermally compressed ambient air 703 from the organpipe intakeaperture to a turbo compressor 710. The intake plenum, which may bedonut shaped, circular, circular compartmented or may be of atwo-dimensional tubular construction, acts both as a diffuser as well asa distribution plenum in event of the use of multiple turbo compressors.The ambient air may enter the intake plenum 704 supersonically or atsonic speed in event of perfect (super) cooling of the shock front.

Compressor 710 of the invention, which is located in the mannerillustrated in FIGS. 7 and 7A of the drawings, compresses the (rarified)intake air at high/hypersonic altitudes to a substantially higherpressure/density (compression ratio=10-15.times. ambient) so as tomitigate the volumetric extent of the sub-cooling heat exchangers andliquefaction plant. Compressor 710 is driven by turbine 711 that is, inturn, driven by superheated nitrogen 751. The compressor may becustom-designed, may be available from conventional suppliers, may be ofa single or multi-stage radial construction, or it may comprise amultistage axial flow compressor.

The previously mentioned turbine 711 is here driven by the separatednitrogen 750 subsequent to being heated via the nosecone and organpipeheat exchangers 701 and 702, respectively.

Considering next the expansion turbine 720, this turbine serves dualpurpose of extracting energy from the compressed intake air 703 andsuper cooling the intake air 703 into the cryogenic zone so as tofacilitate precipitation/separation of liquid oxygen out of the ambientair. Expansion turbine 720 may comprise a radial or axial turbine all orit may comprise a conventional vane motor. The output power of theturbine can be applied to drive the propellant pumps 721 and 722.Additionally, the expansion turbine 720 can also supplement the turbocompressor 710 with a separate/external source (e.g. hydrazine turbine)to drive the propellant pumps. Expansion will be limited to about 5-10times the discharge pressure of the turbocharger compressor 710 so as tomaintain dequate density to traverse the nosecone, organpipe and shroudheat exchangers 701, 702 and intake 706 respectively and additionally toprovide adequate pressure to drive the turbocharger turbine 711.

In operation, pump 721 draws liquid oxygen from either or both theseparator 740 and the liquid oxygen tank 760 depending upon the stage ofthe flight regime and upon the rate of liquid oxygen production by theliquefaction plant and separator 740. Under normal operating conditionspump 721 pumps liquid oxygen thru heat exchanger 750 where it isexpanded while cooling the compressed intake air 703. The expandedoxygen is then ducted by means of a duct 732 to the aerospike engine780.

As indicated in FIGS. 7 and 7A of the drawings, pump 722 draws liquidhydrogen from the liquid hydrogen tank 761 and then pumps the liquidhydrogen thru heat exchanger 751 where it is expanded while sub-coolingthe compressed intake air 703. The expanded hydrogen is then ducted bymeans of duct 733 to the aerospike engine 780.

A first heat exchanger 730, which is driven by the expansion of liquidoxygen that is pumped by means of pump 721 pre-cools the compressedintake air emanating from compressor 710. As indicated in the drawings,the expanded liquid oxygen is then ducted by means of duct 732 to theaerospike rocket motor 780.

A second heat exchanger 731, which is driven by the expansion of liquidhydrogen that is pumped by means of pump 722, sub-cools the compressedintake air emanating from heat exchanger 730. The expanded liquidhydrogen is then ducted by means of duct 733 to the aerospike rocketmotor 780.

In this latest form of the invention, liquid oxygen separator 740separates liquid oxygen from the cryogenically chilled(expanded/saturated) intake air stream 703 emanating from the expansionturbine 720. Separation may be by centrifuge or other by other meanswell known to those skilled in the art. Separation may also be partiallyachieved in the expansion turbine 720.

In FIGS. 7 and 7A of the drawings, numeral 732 identifies the oxygenconduit to the rocket motor, numeral 733 identifies the hydrogen conduitto rocket motor, numeral 740 identifies the Liquid oxygen separator ofthis latest form of the invention, numeral 750 identifies the separatednitrogen conduit leading to the nosecone 701 and numeral 751 identifiesthe superheated nitrogen. More particularly, in operation conduit 750acts a funnel to duct the super-cool nitrogen emanating from separator740 to the nosecone, organpipe and shroud heat exchangers 701, 702 and706 respectively.

As indicated by the flow arrows in FIGS. 7 and 7A, organ pipes 702 allowthe pre-cooled, ambient air 703 to enter the intake apertures and intoplenums 704. From plenums 704 the air flows to turbo compressor 710,into heat exchanger 730 which pre-cool the compressed air emanating fromcompressors 710. From first heat exchanger 730 the air flows to secondheat exchanger 731 which pre-cool the compressed air emanating from heatexchanger 730. From heat exchanger 731 the air flows into expansionturbine 720. As previously mentioned expansion turbine 720 serves thedual purpose of extracting energy from the compressed intake air 703 andsuper cooling the intake air into the cryogenic zone so as to facilitateseparation of liquid oxygen from the ambient air. As best seen in FIG.7, propellant pumps 721 and 722 are operably associated with expansionturbines 720. As previously discussed, pumps 721 function to draw liquidoxygen from both separators 740 as well as liquid oxygen tanks 760. Asindicated in FIG. 7, pump 721 pumps the liquid oxygen through heatexchanger 730 where it is expanded while cooling the compressed intakeair 703. The expanded oxygen is then ducted by means of ducts 732 to theaerospike engine 780. In a similar manner, pump 722 functions to drawliquid hydrogen from hydrogen tank 761 and then pump the liquid hydrogenthrough heat exchanger 731 where it is expanded while sub cooling thecompressed intake air 703. As illustrated in FIG. 7, the expandedhydrogen is then ducted by means of ducts 733 to the aerospike engine780.

Aerospike rocket motor 780, which propels the alternate embodiment ofthe aerospace plane, here comprises a truncated, adaptive expansion ramp781, a hyper expansion ramp 782 and a plume 783. (Plume 783 denotes theexpansion envelope of the aerospike engine 780). Although both the rampsare employed at takeoff and initial fly out, only the hyper expansionramp 782 will be employed through the hypersonic and space sectors ofthe ascent regime. The adaptive aerospike engine is more fully describedin U.S. Pat. No. 6,213,431 issued to the present inventor. Expansionramp 781 is truncated for optimum for take-off. A shorter expansion rampis required for takeoff and during the initial flight due to highatmospheric pressure that counterbalances expansion. The longer hyperexpansion ramp 782 is required at higher altitudes and in space tofacilitate complete expansion with low, or near zero ambient pressure(hence assuring kinetic efficacy at all times). Hyper expansion is morefully discussed in U.S. Pat. No. 6,213,431 issued to the presentinventor. Due to the burning off of propellant truncated ramp 781 isshut down as Mach 10 is approached. In orbit hyper expansion ramp 782will suffice as the propulsive element.

Turning now to FIGS. 7B and 8 of the drawings, yet another embodiment ofthe reversible aerospace plane of the invention is there shown ingenerally designated by the numeral 700A. This latest embodiment of theinvention is similar in many respects to the previously describedembodiments and like numerals are used in FIGS. 7B and 8 to identifylike components.

It is to be noted that in this latest embodiment of the invention, thesuper-cool nitrogen 750 will be heated by means of a heat exchanger, orsuperheat shroud 706 (FIG. 7B) to boost the available turbochargerpower. Being outside the super-cool nosecone envelope, superheat shroud706 is exposed to the full adiabatic compressive/heat impact of theshock front through the hypersonic regime. Accordingly, the nitrogen 751emanating from heat exchanger 706 will be superheated to on the order ofabout 1,000 to about 1,500 F, thereby trebling, or quadrupling thecompression power that may be recaptured in this way.

As before, turbine 711 may be configured either as a turbocharger, oralternatively may comprise an industry standard jet engine turbinespool. In this regard, it is to be noted that singular or multiplespools may be employed in parallel.

Liquid oxygen tank 760 of this latest form of the invention functions asa holding/storage device. In contradistinction to conventional practice,tanking of liquid oxygen will be limited to on the order of 60-80%(pending particularities) of the nominal sector requirement.

The superheat shroud 706 of this latest embodiment of the inventionfunctions as a structural cooling means as well as superheat source toaugment the performance of turbine 711 of the invention which is locateddirectly downstream of the intake plenum 704 and of the organpipe tubes702. The superheat shroud 706 may be finned, warped, dimpled,electroplated, etched or may be of a bimetallic construction.

It is to be observed that because super cooling is limited to the intakeair 703 thru intake plenum 704, the superheat shroud 706 of theinvention is subjected to the (free-flow) adiabatic shock front inexcess of 2,000 deg Fahrenheit. As a consequence, as will be discussedin the paragraphs which follow, because of being ducted through shroud706, the super cool nitrogen 750 will be heated to on the order of 1,500deg Fahrenheit. Accordingly, the ultimate temperature of shroud 706 willbe limited to on the order of 1,000 to 1,500 deg Fahrenheit.

As previously mentioned, the nosecone 701 is configured to act both as ashock cone and heat exchanger and, as in the earlier embodiments of theinvention, is controllably chilled. Chilling of the nosecone is intendedto protect the aerospace plane from the searing hypersonic compressionheat, to mitigate the shock impact by morphing a lower mach number bymeans of super chilling and pre-cooling the ambient/intake air as aninitial step to liquefaction. Chilling the ambient air forcefully will“shock” the compression of the ambient/intake air 703 over the nose coneinto (cool) isothermal in lieu of (hot) adiabatic compression. It isalso to be observed that at a compression ratio of 5 times the heatemanating from a chilled shock front would be approximately half thatwith adiabatic compression. At a compression ratio of 10 times the heatof compression from a chilled shock front would be 8.5 times less. At acompression ratio of 20.times. the heat of compression of a (super)chilled shock front would be 37 times less compared to adiabaticcompression. It is to be noted that the previously identified organpipeconstitutes a continuation of the nosecone cooling process. Moreparticularly, the organpipe tubes are cooled in conjunction withnosecone 701 by means of the super-cool nitrogen 750 emanating fromseparator 740. As previously mentioned, liquid oxygen separator 740separates liquid oxygen from the cryogenically chilled air stream 703emanating from the expansion turbine 720.

Referring now to FIG. 8 of the drawings, the top plan view of thereversible aerospace plane of FIG. 7B is there shown. In this figuredrawing, the para (reentry) and hyperfoil (ascending) wing sections areidentified as 791 and 792, respectively. As before, numerals 780, 781and 782 respectively identify the aerospike engine, the truncated,adaptive expansion ramp and the hyper expansion ramp.

Thus Referring to Tables 1 and 2 which follow, a comparison is theremade between takeoff mass with and without oxygen liquefaction. Withregard to Table 1 is to be observed that liquefaction mitigates thetakeoff weight of the reversible aerospace plane to place a total of80,000 lb in low earth orbit (for example, 25 ft/sec velocity).Conversely, the takeoff weight without liquefaction would be on theorder of 562,118 lb in accordance with the computational example setforth in Table 2. Given that 88% of 562,118 lb would have beenattributed to liquid oxygen (for example, 494,664 lb) in the nominalarrangement, liquefaction would attribute to(562,118−413,521)/494,664=30% savings in tanked liquid oxygen.

TABLE 1 Vu Ep Prr Ag Gg Mu 25000 0.99 9 2.6 2.5 80000 24000 0.99 9 2.54.9 93408 23000 0.99 9 2.4 7.3 108328 22000 0.99 9 2.3 9.5 124788 210000.99 9 2.2 11.6 142791 20000 0.99 9 2.1 13.6 162307 19000 0.99 9 2 15.5183274 18000 0.99 9 1.9 17.3 205592 17000 0.99 9 1.8 19.0 229125 160000.99 9 1.7 20.6 253697 15000 0.99 9 1.6 22.1 279095 14000 0.99 8 1.523.5 305070 13000 0.99 7 1.4 24.8 328202 12000 0.98 6 1.3 26.0 34808311000 0.98 5 1.2 27.0 364678 10000 0.98 4 1.1 28.0 377759 9000 0.98 3 128.9 387472 8000 0.98 2 0.9 29.7 394103 7000 0.95 2 0.8 30.3 398043 60000.95 2 0.7 30.9 401598 5000 0.95 2 0.6 31.4 404630 4000 0.9 2 0.5 31.7407128 3000 0.8 2 0.4 32.0 409192 2000 0.7 3 0.3 32.1 410861 1000 0.6 40.2 32.2 412593 0 0.5 9 0.1 32.2 413521

TABLE 2 Vu Ep Prr Ag Gg Mu 25000 0.99 9 2.6 2.5 80000 24000 0.99 9 2.54.9 93408 23000 0.99 9 2.4 7.3 108328 22000 0.99 9 2.3 9.5 124788 210000.99 9 2.2 11.6 142791 20000 0.99 9 2.1 13.6 162307 19000 0.99 9 2 15.5183274 18000 0.99 9 1.9 17.3 205592 17000 0.99 9 1.8 19.0 229125 160000.99 9 1.7 20.6 253697 15000 0.99 9 1.6 22.1 279095 14000 0.99 9 1.523.5 305070 13000 0.99 9 1.4 24.8 331339 12000 0.98 9 1.3 26.0 35759311000 0.98 9 1.2 27.0 383784 10000 0.98 9 1.1 28.0 409289 9000 0.98 9 128.9 433748 8000 0.98 9 0.9 29.7 456799 7000 0.95 9 0.8 30.3 478090 60000.95 9 0.7 30.9 497919 5000 0.95 9 0.6 31.4 515289 4000 0.9 9 0.5 31.7529912 3000 0.8 9 0.4 32.0 542206 2000 0.7 9 0.3 32.1 552284 1000 0.6 90.2 32.2 559310 0 0.5 9 0.1 32.2 562118

Turning next to FIG. 9, an alternate (two dimensional) form of aerospaceplane of the invention in a parafoil reentry format is there illustratedand generally designated as 990. This latest embodiment comprises ahyperfoil profiled intake 992 that is configured in accordance with thethree dimensional nosecone format. An optional booster rocket 993 isprovided for takeoff. Also provided is a cargo bay, or hydrogen tankagespace that is here identified by the numeral 994. An aerospike rocketengine that can also serve as a reentry heat shield is identified by thenumeral 980.

While the invention has been described with respect to a limited numberof embodiments, those skilled in the art, having benefit of thisdisclosure, will appreciate that other embodiments can be devised whichdo not depart from the scope of the invention as disclosed herein.

I claim:
 1. The reversible aerospace plane comprising: an air intake ata first end of the aerospace plane; a compressor located next to thefirst end of the aerospace plane for compressing an intake air; aturbine located next to said compressor for driving said compressor; afirst heat exchanger disposed in the aerospace plane for cooling thecompressed intake air to produce cooled, compressed intake air; a secondheat exchanger disposed in the aerospace plane for cooling the cooled,compressed intake air; an engine at a second end of the aerospace plane,wherein the aerospace plane is configured to accelerate in a firstorientation and configured to glide and land in a second orientation,wherein the second orientation is substantially a reverse of the firstorientation; wherein one of said first and second heat exchangers isconfigured to use tanked liquid hydrogen as a coolant to condense atleast a portion of an oxygen component of an incident air, and a chillednosecone that is chilled via flashing of the cryogenic propellantwhereby the chilled nosecone instills oxygen liquefaction andcompression of the ambient air stream at hypersonic speed into aone-step process within the confinements of an infinite boundary layeron the peripheral of the cryogenically chilled air stream.
 2. Thereversible aerospace plane of claim 1 whereby the liquefacted ambientair stream is expanded via an expansion turbine that drives thecompressor that compresses the flashed cryogenic propellant to a higherpressure level to a value approximate to the pressure of the expandedliquefacted airstream.
 3. The reversible aerospace plane of claim 1whereby the expanded liquefacted air stream and the compressed cryogenicpropellant is mixed and combusted at high pressure in a combustionchamber at at approximately 1,500 F.
 4. The reversible aerospace planeof claim 1 further comprising a hot combustion gas is expanded via aseries of supersonic nozzles in an adaptive expansion ramp.
 5. Thereversible aerospace plane of claim 1 whereby a hot combustion gasentrains ambient air into the exhaust stream.
 6. The reversibleaerospace plane of claim 1 whereby a combustion gas detonates on impactwith a curved expansion ramp.
 7. The reversible aerospace plane of claim1 whereby the detonated combustion fuel accelerated to a substantivelyhigher speed as result of centrifugal compression of the detonated fuelin a curved expansion ramp.
 8. The reversible aerospace plane of claim 1whereby accelerated combustion fuel is ejected via an aerospike ramp. 9.The reversible aerospace plane of claim 1 that is driven by the reactionforce emanating from an accelerated combustion fuel that is ejected viaan aerospike ramp.
 10. The reversible aerospace plane of claim 1 that isequipped with an induction shroud to induce an additional ambient aircomponent to augment an acceleration thrust vector by momentum exchange.